Method for reprocessing a turbine blade having at least one platform

ABSTRACT

A method for reprocessing a turbine blade having at least one platform which due to the action of corrosion is undersized on at least one lateral platform face is provided. According to the method, the target dimension of the platform is restored by applying material to the at least one lateral platform face such that, after the material application, the platform is oversized and then the platform is given the target dimension by machining the at least one lateral platform face. The material application is carried out with the material of an adhesion promoter layer.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International Application No. PCT/EP2011/053899, filed Mar. 15, 2011 and claims the benefit thereof. The International Application claims the benefits of European Patent Office application No. 10002967.7 EP filed Mar. 19, 2010. All of the applications are incorporated by reference herein in their entirety.

FIELD OF INVENTION

1. Detailed Description of Invention

The present invention relates to a process for refurbishing a turbine blade or vane having at least one platform, wherein the turbine blade or vane can be formed in particular as a gas turbine blade or vane.

2. Background of Invention

In a gas turbine, a liquid or gaseous fuel is burned in a combustion chamber and the hot gases under high pressure which form during the combustion are fed to the turbine, where they transfer momentum to the rotor blades of a turbine with expansion and cooling. In this case, the transfer of momentum to the rotor blades is optimized by means of guide vanes.

Since the hot combustion gases have a strong oxidizing and corrosive action, the turbine blades or vanes, in particular those of the first rows of turbine blades or vanes around which particularly hot combustion gases flow, are produced from superalloys which can withstand high temperatures, and are additionally coated with a thermal barrier coating system, in order to further increase the resistance of the blades or vanes to the oxidizing and corrosive conditions in the hot gas. Such a coating typically comprises a ceramic thermal barrier coating which is bonded to the superalloy material of the blade or vane by means of an adhesion promoter layer. Typical adhesion promoter layers are so-called MCrAlX layers, in which M stands for iron (Fe), cobalt (Co), nickel (Ni) or a combination of these metals. X represents an active element and stands for yttrium (Y) and/or silicon (Si) and/or at least one rare earth element or hafnium (Hf). Such alloys are known, for example, from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

Despite their high resistance to attack by hot gas, corrosion takes place on the blades or vanes by the operational loading and, as a result thereof, by the high-temperature oxidation. This also affects the blade or vane platforms. After a certain operating time, turbine blades or vanes are therefore subjected to a refurbishing process, in which the coating is removed, sites damaged by corrosion are repaired and the blades or vanes are then recoated, in order to prepare them for renewed use in a gas turbine.

Particularly in the case of the turbine stages through which the hottest combustion gases flow, typically in the case of the first two stages, the corrosive attack can lead to an undersize on the platform side faces, however.

EP 1 808 266 A2 proposes removing platform regions damaged by corrosion in the region of the trailing edge of the turbine blade or vane and then rebuilding the removed region by build-up welding and subsequent grinding to the correct dimension. Although in principle the undersized side faces of platforms can also be built up again in this way, build-up welding on superalloy materials is difficult. In particular, undesirable structural properties in the superalloy material which weaken the material can arise on account of the introduction of heat.

SUMMARY OF INVENTION

It is an object of the present invention, therefore, to provide an advantageous process for refurbishing a turbine blade or vane having at least one platform.

This object is achieved by a process for refurbishing a turbine blade or vane as claimed in the claims The dependent claims contain advantageous configurations of the invention.

In the process according to the invention for refurbishing a turbine blade or vane having at least one platform which, on account of corrosive attack, is undersized on at least one platform side face, the desired dimension of the platform is restored by applying material to the at least one platform side face in such a manner that, after the material application, the platform is oversized, and then the platform is given the desired dimension by machining the at least one platform side face. According to the invention, the material application is effected with the material of an adhesion promoter layer. This material can be, in particular, an MCrAlX material.

Compared with build-up welding, the invention has the advantage that the application of an adhesion promoter material, in particular of MCrAlX material, does not entail such a high introduction of heat into the superalloy material as would be the case, for example, in the case of build-up welding. As the side face is being built up, the microstructure of the superalloy material is therefore disturbed to a lesser extent by means of the adhesion promoter material than in the case of application by means of build-up welding. In addition, the material application can be integrated in the process of recoating the turbine blade or vane, since an adhesion promoter layer is also applied when reapplying a thermal barrier coating system. The process according to the invention therefore makes it possible, in a cost-effective and gentle manner, to restore the desired dimension of platform side faces in operationally stressed turbine blades or vanes, as a result of which the proportion of rejects of operationally stressed turbine blades or vanes can be reduced.

In the context of the process according to the invention, the material application can be effected in particular by means of the repeated application of adhesion promoter material. A material application of at least 10 μm, preferably at least 30 μm, can be effected in this case in particular each time the application of adhesion promoter material is repeated.

For better bonding of the adhesion promoter material to the superalloy material, a bonding heat treatment can take place after the application of the adhesion promoter material.

In the context of the process according to the invention, the material application and the machining can also be effected in particular on two oppositely positioned platform faces of a blade or vane platform. Specifically, it is often the case that undersized regions caused by corrosion arise on two oppositely positioned sides of blade or vane platforms at the same time.

Typically, the turbine blade or vane has a central axis. It is therefore advantageous if the machining of the platform side faces to give the platform the desired dimension again after the application of material is effected with respect to the central axis. It is thereby possible to ensure that not only the platform width but also the distance between the platform side faces and the main blade or vane part of the turbine are given the desired dimension again. To this end, by way of example, the current dimension of the platform can be gathered by scanning at least five measurement points on the opposite platform side faces. The required material removal for the machining is then determined from the current dimension. In this context, the two opposite platform faces can in particular be the platform side faces which are located on the pressure side and suction side in relation to a main blade or vane part with a pressure side and a suction side. During operation of a gas turbine, these side faces are typically exposed to the hot gas oxidation and the resulting corrosion to a greater extent than the platform side faces located on the inflow side and on the outflow side.

In the context of the process according to the invention, the machining can be realized in particular by face grinding.

If the adhesion promoter material is applied within the scope of renewing a thermal barrier coating system of the turbine blade or vane, the process can comprise the removal of layers from the turbine blade or vane before the thermal barrier coating system is renewed. In addition, activation blasting can be effected after the layers have been removed and before the thermal barrier coating system is renewed. The activation blasting would then also include in particular the platform side faces onto which material is to be applied. In such activation blasting, the surfaces are irradiated by means of a blasting agent, for example by means of aluminum oxide (Al₂O₃), as a result of which the surface is roughened, which improves the adhesion of the adhesion promoter material to be applied.

The adhesion promoter material can be applied using a thermal spraying process, for example plasma spraying, flame spraying, etc. Such processes are known as possible processes for applying adhesion promoter layers and can therefore also be used in a readily manageable manner for the application of material to undersized platform side faces.

BRIEF DESCRIPTION OF THE DRAWINGS

Further features, advantages and properties of the present invention will become apparent from the following description of exemplary embodiments with reference to the accompanying figures.

FIG. 1 shows a schematic illustration of a gas turbine in a partial longitudinal section.

FIG. 2 shows an example of a combustion chamber of a gas turbine in a partially sectioned, perspective illustration.

FIG. 3 shows an example of a turbine blade or vane in a perspective illustration.

FIG. 4 shows a schematic plan view of a turbine blade or vane which is undersized as a result of corrosion on side faces of the platform.

FIG. 5 shows the turbine blade or vane shown in FIG. 4 during the application of adhesion promoter material.

FIG. 6 shows the turbine blade or vane shown in FIG. 4 during the grinding of the applied adhesion promoter material to the desired dimension.

FIG. 1 shows, by way of example, a partial longitudinal section through a gas turbine 100.

In the interior, the gas turbine 100 has a rotor 103 with a shaft 101 which is mounted such that it can rotate about an axis of rotation 102 and is also referred to as the turbine rotor.

An intake housing 104, a compressor 105, a, for example, toroidal combustion chamber 110, in particular an annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust-gas housing 109 follow one another along the rotor 103.

The annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111, where, by way of example, four successive turbine stages 112 form the turbine 108.

Each turbine stage 112 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium 113, in the hot-gas passage 111 a row of guide vanes 115 is followed by a row 125 formed from rotor blades 120.

The guide vanes 130 are secured to an inner housing 138 of a stator 143, whereas the rotor blades 120 of a row 125 are fitted to the rotor 103 for example by means of a turbine disk 133.

A generator (not shown) is coupled to the rotor 103.

While the gas turbine 100 is operating, the compressor 105 sucks in air 135 through the intake housing 104 and compresses it. The compressed air provided at the turbine-side end of the compressor 105 is passed to the burners 107, where it is mixed with a fuel. The mix is then burnt in the combustion chamber 110, forming the working medium 113. From there, the working medium 113 flows along the hot-gas passage 111 past the guide vanes 130 and the rotor blades 120. The working medium 113 is expanded at the rotor blades 120, transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the latter in turn drives the generator coupled to it.

While the gas turbine 100 is operating, the components which are exposed to the hot working medium 113 are subject to thermal stresses. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, as seen in the direction of flow of the working medium 113, together with the heat shield elements which line the annular combustion chamber 110, are subject to the highest thermal stresses.

To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant.

Substrates of the components may likewise have a directional structure, i.e. they are in single-crystal form (SX structure) or have only longitudinally oriented grains (DS structure).

By way of example, iron-based, nickel-based or cobalt-based superalloys are used as material for the components, in particular for the turbine blade or vane 120, 130 and components of the combustion chamber 110.

Superalloys of this type are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blades or vanes 120, 130 may likewise have coatings protecting against corrosion (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon, scandium (Sc) and/or at least one rare earth element, or hafnium). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

It is also possible for a thermal barrier coating to be present on the MCrAlX, consisting for example of ZrO₂, Y₂O₃—ZrO₂, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).

The guide vane 130 has a guide vane root (not shown here), which faces the inner housing 138 of the turbine 108, and a guide vane head which is at the opposite end from the guide vane root. The guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143.

FIG. 2 shows a combustion chamber 110 of a gas turbine.

The combustion chamber 110 is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity of burners 107, which generate flames 156, arranged circumferentially around an axis of rotation 102 open out into a common combustion chamber space 154. For this purpose, the combustion chamber 110 overall is of annular configuration positioned around the axis of rotation 102.

To achieve a relatively high efficiency, the combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C. To allow a relatively long service life even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed from heat shield elements 155.

On the working medium side, each heat shield element 155 made from an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) or is made from material that is able to withstand high temperatures (solid ceramic bricks).

These protective layers may be similar to the turbine blades or vanes, i.e. for example MCrAlX: M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

It is also possible for a, for example ceramic, thermal barrier coating to be present on the MCrAlX, consisting for example of ZrO₂, Y₂O₃—ZrO₂, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).

Other coating processes are possible, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks.

Refurbishment means that after they have been used, protective layers may have to be removed from heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the heat shield element 155 are also repaired. This is followed by recoating of the heat shield elements 155, after which the heat shield elements 155 can be reused.

Moreover, a cooling system may be provided for the heat shield elements 155 and/or their holding elements, on account of the high temperatures in the interior of the combustion chamber 110. The heat shield elements 155 are then, for example, hollow and may also have cooling holes (not shown) opening out into the combustion chamber space 154.

FIG. 3 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor.

The blade or vane 120, 130 has, in succession along the longitudinal axis 121, a securing region 400, an adjoining blade or vane platform 403 and a main blade or vane part 406 and a blade or vane tip 415.

As a guide vane 130, the vane 130 may have a further platform (not shown) at its vane tip 415.

A blade or vane root 183, which is used to secure the rotor blades 120, 130 to a shaft or a disk (not shown), is formed in the securing region 400.

The blade or vane root 183 is designed, for example, in hammerhead form. Other configurations, such as a fir-tree or dovetail root, are possible.

The blade or vane 120, 130 has a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade or vane part 406.

In the case of conventional blades or vanes 120, 130, by way of example solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade or vane 120, 130.

Superalloys of this type are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.

The blade or vane 120, 130 may in this case be produced by a casting process, by means of directional solidification, by a forging process, by a milling process or combinations thereof.

Workpieces with a single-crystal structure or structures are used as components for machines which, in operation, are exposed to high mechanical, thermal and/or chemical stresses.

Single-crystal workpieces of this type are produced, for example, by directional solidification from the melt. This involves casting processes in which the liquid metallic alloy solidifies to form the single-crystal structure, i.e. the single-crystal workpiece, or solidifies directionally.

In this case, dendritic crystals are oriented along the direction of heat flow and form either a columnar crystalline grain structure (i.e. grains which run over the entire length of the workpiece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of one single crystal. In these processes, a transition to globular (polycrystalline) solidification needs to be avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries, which negate the favorable properties of the directionally solidified or single-crystal component.

Where the text refers in general terms to directionally solidified microstructures, this is to be understood as meaning both single crystals, which do not have any grain boundaries or at most have small-angle grain boundaries, and columnar crystal structures, which do have grain boundaries running in the longitudinal direction but do not have any transverse grain boundaries. This second form of crystalline structures is also described as directionally solidified microstructures (directionally solidified structures).

Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1.

The blades or vanes 120, 130 may likewise have coatings protecting against corrosion or oxidation e.g. (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.

The density is preferably 95% of the theoretical density.

A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).

The layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protective coatings, it is also preferable to use nickel-based protective layers, such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re.

It is also possible for a thermal barrier coating, which is preferably the outermost layer, to be present on the MCrAlX, consisting for example of ZrO₂, Y₂O₃—ZrO₂, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

The thermal barrier coating covers the entire MCrAlX layer.

Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).

Other coating processes are possible, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks. The thermal barrier coating is therefore preferably more porous than the MCrAlX layer.

Refurbishment means that after they have been used, protective layers may have to be removed from components 120, 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the component 120, 130 are also repaired. This is followed by recoating of the component 120, 130, after which the component 120, 130 can be reused.

The blade or vane 120, 130 may be hollow or solid in form. If the blade or vane 120, 130 is to be cooled, it is hollow and may also have film-cooling holes 418 (indicated by dashed lines).

FIG. 4 shows a schematic illustration of a corroded turbine blade or vane. The illustration shows the main blade or vane part 1 with a suction side 3 and pressure side 5 and also the blade or vane platform 7. On account of corrosive attacks, which originate from the operational loading and the resulting high-temperature oxidation, said blade or vane platform is undersized both on the suction-side platform side face 9 and on the pressure-side platform side face 11. The desired dimension of the platform 7 is shown by dashed lines in the figure. At this point, it should be pointed out that the undersize induced by corrosion is shown in exaggerated form in order to enhance the clarity of the drawings. Corrosion in regions of the turbine blade or vane other than the side faces 9, 11 of the blade or vane platform 7 is typically also present, but is not shown in the figures so as not to unnecessarily complicate the drawings. The turbine blade or vane which has been corroded by the harsh ambient conditions which prevail during gas turbine operation is subjected according to the invention to refurbishment, in which the desired dimension of the blade or vane platform, in particular at the corroded side faces 9, 11, is restored.

In the present exemplary embodiment, the restoration of the desired dimension is in part integrated in the procedure for applying a new thermal barrier coating system to the turbine blade or vane. To this end, the old coating is firstly removed from the turbine blade or vane, for example by means of suitable solutions and/or suitable blasting processes, and the blade or vane is then cleaned in order to remove possible oxidation residues. This is followed by activation blasting, for example by means of aluminum oxide particles (Al₂O₃, corundum), by means of which the surface is roughened. The thus prepared turbine blade or vane is then introduced into a coating apparatus in order to then coat the main blade or vane part 1 and the face 13 of the blade or vane platform 7 which faces toward the hot loading path with the thermal barrier coating system. In this respect, the side faces of the blade or vane platform 7 are usually masked or obscured in the prior art, since they are not intended to be provided with a thermal barrier coating system. Within the context of the process according to the invention, however, the suction-side platform side face 9 and also the pressure-side platform side face 11 are not masked, or obscured, but instead are left free. It is thereby possible to apply coating material to these two faces.

The coating process firstly involves the application of an adhesion promoter layer, which in the present example is in the form of an MCrAlX layer. The adhesion promoter layer is applied by means of a thermal spraying process, for example by means of plasma spraying or flame spraying. The MCrAlX material is applied not only to the main blade or vane part 1 and the top side 13 of the blade or vane platform 7, but also to the suction-side platform side face 9 and the pressure-side platform side face 11.

The spraying process is indicated in FIG. 5 by a schematically shown spray nozzle 15. Unlike in the other regions, however, a plurality of layers of the MCrAlX material are applied to the suction-side platform side face 9 and the pressure-side platform side face 11, with each layer having a minimum thickness of 10 μm, preferably of 30 μm. The layered application of MCrAlX material 12 to the platform side faces 9, 11 is effected until the desired dimension of the platform 7, as indicated by dashed lines in FIGS. 4 to 6, is exceeded. This state is shown in FIG. 5 for the suction-side platform side face 9, whereas the layered spraying-on of the MCrAlX material 12 by means of the spray nozzle 15 is shown for the pressure-side platform side face 11.

Once so many layers of MCrAlX material 12 have been applied to both platform side faces 9, 11 that the desired dimension is exceeded for both, a preferred bonding heat treatment which improves the bonding of the applied MCrAlX material 12 to the superalloy material of the turbine blade or vane is carried out. This is followed by the application of a thermal barrier coating, for example a zirconium oxide layer (ZrO₂), the structure of which is stabilized at least partially by yttrium oxide (Y₂O₃), to the MCrAlX layer. The thermal barrier coating is applied in particular to the main blade or vane part 1 and the surface 13 of the platform 7. Within the context of the invention, however, there is also no disruption if the thermal barrier coating is also applied to the MCrAlX material 12 applied to the platform side faces 9, 11. This will often also be unavoidable since, as mentioned above, the platform side faces 9, 11 are not covered or obscured. The thermal barrier coating can likewise be applied by means of a thermal spraying process. Alternatively, however, it is also possible to produce the thermal barrier coating by vapor deposition.

Once the thermal barrier coating has been applied to the turbine blade or vane, the latter is removed from the coating apparatus and clamped in an apparatus for machining, in which, in the present exemplary embodiment, the now oversized suction-side and pressure-side platform side faces are then machined by means of grinding. The turbine blade or vane is clamped in such a way here that the central axis A of the turbine blade or vane is the same as the central axis of the clamping apparatus. The turbine blade or vane is clamped for grinding such that the turbine blade or vane can be freely positioned in space with respect to its central axis A, i.e. can be rotated through 360°.

The current width b of the blade or vane platform 7 is then captured by sampling at least 5 measurement points for each platform side face 9, 11 on the pressure side and suction side. A computer program is then used to calculate the material removal which is required to give the blade or vane platform 7 having the applied MCrAlX material the desired dimension. The machining removal calculated for the suction-side platform side face 9 and the pressure-side platform side face 11 is in this case based on the blade or vane central axis A. The ascertained material to be removed is then machined off by face grinding by means of a grinding apparatus, which is shown in greatly schematic form in FIG. 6 under the reference numeral 17. Before zero grinding, the grinding disk is removed and the truing amount is compensated by the program, based on the disk diameter of the grinding apparatus. After the grinding process has been completed, the width b of the blade or vane platform 7 has the desired dimension again.

Once the turbine blade or vane has been recoated and the platform side faces 9, 11 have been given the desired dimension again, the turbine blade or vane can be reinstalled into a gas turbine for further operation.

The present invention was described on the basis of a specific exemplary embodiment for the purposes of explanation. Deviations from this exemplary embodiment are possible, however. By way of example, it is thus not absolutely necessary to carry out the application of the adhesion promoter material in the course of a process for reapplying a thermal barrier coating system to a turbine blade or vane. Instead, adhesion promoter material can be applied to the platform side faces in an independent process. Similarly, it is not absolutely necessary to apply material to the platform side faces both on the suction side and on the pressure side. Even though both sides as a rule are undersized on account of corrosion, it may also be the case that only one side is undersized. In this case, the application of material to the undersized side is sufficient. Nevertheless, it may be desirable, however, to also apply material to the side which is not considerably undersized, and then to grind said material, in order to achieve an improved surface structure of the platform side face. 

1-12. (canceled)
 13. A process for refurbishing a turbine blade or vane having at least one platform which, on account of corrosive attack, is undersized on a platform side face, comprising: restoring the desired dimension of the platform, the restoring comprising: applying material to the platform side face in such a manner that, after the material application, the platform is oversized, and machining the platform side face in order the to give the desired dimension to the platform, wherein the material application is effected with the material of an adhesion promoter layer, wherein the adhesion promoter material is applied within the scope of renewing a thermal barrier coating system of the turbine blade or vane which comprises an adhesion promoter layer and a thermal barrier coating, and wherein the material of the adhesion promoter layer is an MCrAlX material.
 14. The process as claimed in claim 13, wherein the material application is effected by means of the repeated application of the adhesion promoter material.
 15. The process as claimed in claim 13, wherein the material application of at least 10 um is used each time the application of adhesion promoter material is repeated.
 16. The process as claimed in claim 13, further comprising using a bonding heat treatment after the application of adhesion promoter material.
 17. The process as claimed in claim 13, wherein the material application and the machining are effected on two opposite platform side faces.
 18. The process as claimed in claim 17, wherein the turbine blade or vane includes a central axis and the machining of the platform side face to give the platform the desired dimension is effected with respect to the central axis.
 19. The process as claimed in claim 18, wherein the current dimension of the platform is gathered by scanning at least five measurement points on the opposite platform side faces and the required material removal for the machining is determined from the current dimension.
 20. The process as claimed in claim 17, wherein the turbine blade or vane includes a main blade or vane part with a pressure side and a suction side and the opposite platform side faces are the platform side faces located on the pressure side and suction side in relation to the main blade or vane part.
 21. The process as claimed in claim 13, the machining is realized by face grinding.
 22. The process as claimed in claim 21, wherein a plurality of layers are removed from the turbine blade or vane before the thermal barrier coating system is renewed.
 23. The process as claimed in claim 22, wherein activation blasting is effected after the plurality of layers have been removed and before the thermal barrier coating system is renewed.
 24. The process as claimed in claim 13, wherein the adhesion promoter material is applied by means of a thermal spraying process. 